Gas turbine engines, such as turbofan gas turbine engines, may be used to power various types of vehicles and systems, such as aircraft. Typically, these engines include turbines that rotate at a high speed when blades (or airfoils) extending therefrom are impinged by high-energy compressed air. Consequently, the blades are subjected to high heat and stress loadings which, over time, may reduce their structural integrity.
To improve blade structural integrity, a blade cooling scheme is typically incorporated into the turbines. The blade cooling scheme is included to maintain the blade temperatures within acceptable limits. In some cases, the blade cooling scheme directs cooling air through an internal cooling circuit formed in the blade. The internal cooling circuit may include a simple channel extending through a length of the blade or may consist of a series of connected, serpentine cooling passages, which incorporate raised or depressed structures therein. The serpentine cooling passages increase the cooling effectiveness by extending the length of the air flow path. In this regard, the blade may have multiple internal walls that form the intricate cooling passages through which the cooling air flows.
As the desire for increased engine efficiency continues to rise, engine components are increasingly being subjected to higher and higher operating temperatures. For example, newer engine designs may employ operating temperatures that are over 1100° C. However, current engine components, such as the blades, may not be adequately designed to withstand such temperatures over time. Hence, designs for improving cooling of the blades may be desired.
Turbine blade tips (at the extreme outer radial region) are difficult to cool due to geometry, manufacturing constraints, and the high velocity air that migrates from the pressure side of the airfoil to the suction side via the gap between the rotor tip and the turbine shroud. The trailing edge of the blade tip is particularly difficult to cool in a manner that does not detrimentally affect the turbine performance or introduce risk.
Hence, there is an unmet need in the art for a turbine blade having a cooling system that is capable of cooling the blade tip in high-temperature operating environments. The present disclosure addresses at least this need.